Joints between a composite skin and a load-bearing component and methods of forming same

ABSTRACT

A method of forming a direct bearing joint includes providing a load bearing structure including a first structural feature having an arcuate shape and providing a composite skin including a second structural feature also having an arcuate shape. The method also includes coupling the first structural feature to the second structural feature such that the first and second structural features are mated against each other to facilitate distributing compressive loads induced into the load bearing structure into the composite skin.

BACKGROUND

The implementations described herein relate generally to forming a jointbetween two or more mechanical components, and, more specifically, toforming a direct bearing joint between a load-bearing component andcomposite component.

At least some known aircraft are designed and manufactured using largeamounts of composite materials. For example, composite materials areused in an aircraft to decrease the weight of the aircraft. Decreasingthe overall weight may improve performance features, such as, forexample, payload capacities and fuel efficiencies. At least some knownaircraft include a fuselage structure including a thin load-bearingcomposite skin supported by circumferential bulkheads designed totransfer shear stresses and to retain the shape of the fuselage. Majorload carrying metal components, such as wings spars, the keel beam, andwheel well longerons, for example, are coupled to both the bulkheads andthe composite skin to distribute concentrated loads to the compositeskin.

During at least some known metal or composite aircraft construction,bulkheads, major load bearing components, and composite skin are joinedtogether via metal fasteners. As the amount of load to be distributedincreases, the number and size of fasteners required to join thecomponents together increases. However, because fasteners have a minimumspacing allotment from each other and from the skin edge, increasing thesize and number of fasteners also increases the overall size and weightof the joint required to join the components. In some cases, the jointmay become impractical because of space limitations if for example, thelarger joint may interfere with other systems of the aircraft.Furthermore, in at least some known aircraft, a gap may be definedbetween the skin and the structure transferring load to the skin, suchthat the load path extends only through the fasteners that couple theskin to the structure. In such cases, if one or more of the fastenersfail, the load would be distributed to the remaining fasteners, thusincreasing the risk of their failure as well.

At least one known construction solution is to distribute theconcentrated load over multiple locations along the aircraft such thateach location has less load to bear. However, such a construction methodstill results in a net increase of fasteners that are distributed acrossadditional joints. Moreover, in similar cases the additional joints mayalso interfere with other aircraft systems, such as the landing gear. Assuch, increasing the quantity of fasteners or the size of the fastenersto join various aircraft components generally result in a significantincrease in weight, complexity, and cost penalties for larger loads.

BRIEF DESCRIPTION

In one aspect, a method of forming a direct bearing joint is provided.The method includes providing a load bearing structure including a firststructural feature having an arcuate shape and providing a compositeskin including a second structural feature also having an arcuate shape.The method also includes coupling the first structural feature to thesecond structural feature such that the first and second structuralfeatures are mated against each other to facilitate distributingcompressive loads induced into the load bearing structure into thecomposite skin.

In another aspect, a direct bearing joint is provided. The directbearing includes a load bearing structure including a first structuralfeature having an arcuate shape and a composite skin that is coupled tothe load bearing structure. The composite skin includes a secondstructural feature that also has an arcuate shape. The second structuralfeature mates against the first structural feature to facilitatedistributing compressive loads induced into the load bearing structureinto the composite skin.

In yet another aspect, an aircraft comprising a direct bearing joint isprovided. The aircraft includes a load bearing structure including afirst structural feature having an arcuate shape and a composite skinthat is coupled to the load bearing structure. The composite skinincludes a second structural feature that also has an arcuate shape. Thesecond structural feature mates against the first structural feature tofacilitate distributing compressive loads induced into the load bearingstructure into the composite skin.

BRIEF DESCRIPTION OF THE DRAWINGS

FIG. 1 is a flow diagram of an exemplary aircraft production and servicemethodology;

FIG. 2 is a block diagram of an exemplary aircraft;

FIG. 3 is a schematic diagram of an exemplary joint environment of anaircraft;

FIG. 4 is a side view of an exemplary aircraft fuselage illustrating amain landing gear bay;

FIG. 5 is an enlarged view of a portion of the aircraft fuselage shownin FIG. 4 and illustrating an exemplary direct bearing joint;

FIG. 6 is a cross-sectional view of the direct bearing joint shown inFIG. 4 and taken at location 6-6;

FIG. 7 is a cross-sectional view of the direct bearing joint shown inFIG. 4 and taken at location 7-7;

FIG. 8 is a cross-sectional view of an alternative direct bearing jointthat may be used with the aircraft fuselage shown in FIG. 4 and taken atlocation 8-8;

FIG. 9 is an alternative direct bearing joint that may be used with theaircraft fuselage shown in FIG. 4 and taken at location 9-9; and

FIG. 10 is a cross-sectional view of the alternative direct bearingjoint shown in FIG. 9 and taken at location 10-10.

DETAILED DESCRIPTION

The implementations described herein relate to direct bearing joints andmethods of forming the same. More specifically, in the exemplaryenablement, the direct bearing joint is formed as a continuous interfacebetween a composite skin and a metallic load bearing structure. In theexemplary implementation, the composite skin includes an arcuate cutoutthat mates against a correspondingly-shaped feature on the load bearingstructure at a location where a compressive point load is applied to thestructure. The methods described herein facilitate distributing largecompressive point loads by directly transferring at least a portion ofthe load into the skin, such that a first load path is formed at thedirect bearing joint between the skin and the structure and such that asecond load path is defined through a plurality of fasteners that couplethe skin and structure together. Multiple loads paths reduce the numberand size of the fasteners required to transfer loads into the compositeskin, and thus reduce the overall weight and cost of the structureincluding the direct bearing joint. The plurality of fasteners alsoserve to transfer tension and shear loads that cannot be transferred viadirect bearing.

Referring now to FIG. 1, implementations of the disclosure may bedescribed in the context of an aircraft manufacturing and service method100 and via an aircraft 102. During pre-production, includingspecification and design 104 data of aircraft 102 may be used during themanufacturing process and other materials associated with the airframemay be procured 106. During production, component and subassemblymanufacturing 108 and system integration 110 of aircraft 102 occurs,prior to aircraft 102 entering its certification and delivery process112. Upon successful satisfaction and completion of airframecertification, aircraft 102 may be placed in service 114. While inservice by a customer, aircraft 102 is scheduled for periodic, routine,and scheduled maintenance and service 116, including any modification,reconfiguration, and/or refurbishment, for example. In alternativeimplementations, manufacturing and service method 100 may be implementedvia vehicles other than an aircraft.

Each portion and process associated with aircraft manufacturing and/orservice 100 may be performed or completed by a system integrator, athird party, and/or an operator (e.g., a customer). For the purposes ofthis description, a system integrator may include without limitation anynumber of aircraft manufacturers and major-system subcontractors; athird party may include without limitation any number of venders,subcontractors, and suppliers; and an operator may be an airline,leasing company, military entity, service organization, and so on.

As shown in FIG. 2, an aircraft 102 produced via method 100 may includean airframe 118 including a plurality of systems 120 and an interior122. Exemplary high-level systems 120 include one or more of apropulsion system 124, an electrical system 126, a hydraulic system 128,and/or an environmental system 130. Any number of other systems may beincluded. In such examples, airframe 118 may be formed from structures132 that may be coupled together with joints 134. Moreover, structures132 may include, for example, without limitation, skin panels, wingboxes, stabilizers, longerons, keel beam, spars, ribs, and/or othersuitable types of structures for airframe 118. Although an aerospaceexample is illustrated via aircraft 102, in alternative implementations,the technology may be used with non-aviation industries, such as theautomotive and/or marine industries.

Apparatus and methods embodied herein may be employed during any one ormore of the stages of method 100. For example, components orsubassemblies corresponding to component production process 108 may befabricated or manufactured in a manner similar to components orsubassemblies produced while aircraft 102 is in service. Also, one ormore apparatus implementations, method implementations, or a combinationthereof may be utilized during the production stages 108 and 110, forexample, by substantially expediting assembly of, and/or reducing thecost of assembly of aircraft 102. Similarly, one or more of apparatusimplementations, method implementations, or a combination thereof may beutilized while aircraft 102 is being serviced or maintained, forexample, during scheduled maintenance and service 116.

As used herein, the term “aircraft” may include, but is not limited toonly including, airplanes, unmanned aerial vehicles (UAVs), gliders,helicopters, and/or any other object that travels through airspace.Further, in an alternative implementation, the aircraft manufacturingand service method described herein may be used in any manufacturingand/or service operation.

FIG. 3 is a schematic diagram of an exemplary joint environment 200 thatmay be used to implement a joint for an aircraft, such as, for example,joints 134 (shown in FIG. 2) of aircraft 102 (shown in FIGS. 1 and 2).In the exemplary implementation, a direct bearing joint 202 may beformed between a first aircraft structure 204 and a second aircraftstructure 206. Joint 202 is formed where at least first structure 204and second structure 206 are coupled together using a plurality offasteners 208. In the exemplary implementation, first aircraft structure204 is a skin panel 210 used in a wing (not shown) or the fuselage (notshown in FIG. 3) of airframe 118 (shown in FIG. 2) and second aircraftstructure 206 is a major load bearing component of airframe 118, suchas, but not limited to, at least one of a wheel well longeron (not shownin FIG. 3), a keel beam (not shown in FIG. 3), and a wing rear sparframe (not shown). In the exemplary implementation, second aircraftstructure 206 is fabricated from a metallic material, such as, but notlimited to titanium, aluminum, steel, or any combination thereof.Alternatively, second aircraft structure 206 may be fabricated from anysuitable metal or metal alloy that facilitates operation of joint 202 asdescribed herein.

Skin panel 210 includes a plurality of composite material layers thatare laminated together to form a composite skin. As used herein, a“laminated” object refers to an object fabricated with laminates, andtypically includes multiple layers or plies of composite that includefibers in a resin, or metallic layers. The layers of composite materialmay be formed in any manner desired for constructing skin panel 210. Forexample, different layers may be associated with different angles ororientations with respect to other layers, depending on the particularimplementation. Further, resin and other materials used in the compositematerial also may vary, depending on the particular implantation.

Furthermore, as used herein, “composite” means engineered materialsfabricated from two or more constituent materials. In one exemplaryimplementation, the composite material is a carbon fiber reinforcedpolymer (CFRP) composite, that includes carbon fiber embedded in amatrix or resin, especially epoxy matrices, thermosetting orthermoplastic resins. CFRP composites have an advantageousstrength/weight ratio given that it is a material that is generallylightweight, but possess strong structural properties. Alternatively,other materials and other composite materials may be used, includingthose containing, but not limited to, fiberglass, ceramics, and/or otherelements.

In locations where joints are formed between composite materials andmetallic structures, some joints may be required to carry higher loadsthan other joints. Generally, direct bearing joint 202 may be formed bycoupling first aircraft structure 204 to second aircraft structure 206.More specifically, direct bearing joint 202 is formed by couplingcomposite fuselage skin 210 to at least one of the wing rear spar frame(not shown), the keel beam (not shown in FIG. 3), and/or the wheel welllongeron (not shown in FIG. 3). In the exemplary implementation, directbearing joint 202 is positioned along airframe 118 such that joint 202receives a significant concentrated compressive load, as shown by arrow212. Load 212 is direct into second aircraft structure 206 and then issubsequently transferred into skin panel 210 through direct bearingjoint 202 for distribution throughout the fuselage. Direct bearing joint202 includes a first feature 214 that is cutout of an edge 216 of skin210 and a second feature 218 extending from second structure 206. Firstfeature 214 includes a shape that is complementary to the shape ofsecond feature 218 such that features 214 and 218 are in intimatecontact along direct bearing joint 202. In such a configuration, thedirect contact between skin 210 and second structure 206 provides fordistribution of concentrated compression load 212 through skin 210 toreduce stress concentrations and enable reaction of non-vertical loadcomponents.

Furthermore, direct bearing joint 202 serves as a primary load path forload 212 distribution into skin 210. The plurality of fasteners 208serves a secondary, or fail-safe, load path. The use of direct bearingjoint 202 enables load 212 to be distributed into skin 210 using fewerand smaller fasteners 208 than would be required if fasteners 208 servedas the primary load path. Fewer fasteners 208 facilitates reducing theoverall weight of the aircraft and also reducing costs through less rawmaterials and simpler assembly. In the exemplary implementation, directbearing joint 202 is substantially arch-shaped to facilitate evenlydistributing load 212 throughout skin 210. Alternatively, direct bearingjoint 202 may have any shape, such as, but not limited to, a step shapeor an angular shape, that facilitates operation of joint 202 asdescribed herein.

In certain implementations, direct bearing joint 202 may be used inother platforms other than aircraft 102. For example, direct bearingjoint 202 may be used in a platforms including, but not limited to, amobile platform, a stationary platform, a land-based structure, anaquatic-based structure, a space-based structure, and/or some othersuitable object. More specifically, direct bearing joint 202 may be usedwithin, for example, without limitation, a submarine, a bus, a personnelcarrier, a tank, a train, an automobile, a spacecraft, a space station,a satellite, a surface ship, a power plant, a dam, a bridge, amanufacturing facility, and/or a building.

FIGS. 4-8 illustrate an exemplary direct bearing joint 202 formedbetween first structure 204, such as composite skin 210 (shown in FIG.3) and second structure 206, such as the wheel well longeron, with theunderstanding that such a joint 202 is not limited to being formedbetween these two components, but may also be formed between skin 210and a keel beam and/or between skin 210 and a wing rear spar.

Referring now to FIG. 4, a side view of an exemplary aircraft 300 ispresented. Aircraft 300 is substantially similar to aircraft 102 (shownin FIGS. 1 and 2) and may be manufactured and serviced using method 100(shown in FIG. 1). In the exemplary implementation, aircraft 300includes a fuselage 302 that includes a composite skin 304, which issubstantially similar to skin 210, a plurality of frames 306, and atleast one stringer or longeron, such as a wheel well longeron 308,wherein skin 304, frames 306, and wheel well longeron 308 are coupledtogether by at least a plurality of fasteners 310. Frames 306 are spacedlongitudinally through fuselage 302 along a longitudinal axis 312, andare configured to support skin 304 to provide support to fuselagestringers, and/or to introduce discrete loads into skin 304.

In the exemplary implementation, a direct bearing joint 314 is formedbetween composite first aircraft structure 204 and metallic secondaircraft structure 206 that are coupled together using a plurality offasteners, such as fasteners 208. More specifically, in theimplementation, first aircraft structure 204 is skin 304, secondaircraft structure 206 is wheel well longeron 308, and fasteners 208 arefasteners 310. Fuselage 302 also includes a wing profile section 316that includes a front spar fitting 318 and a rear spar fitting 320. Amain landing gear wheel well 322 is positioned aft of rear spar fitting320 for use in stowing the main landing gear (not shown) when retractedand is adjacent to direct bearing joint 314. In the exemplaryimplementation, fuselage 302 also includes a forward main landing gearjoint 324 that is positioned adjacent to direct bearing joint 314 and anaft main landing gear joint 326 that is positioned aft of forward mainlanding gear joint 324. Joints 324 and 326 couple the main landing gearto fuselage 302. Alternatively, the main landing gear of aircraft 300may be coupled to fuselage 302 by any number of joints in any positionalong fuselage 302 that facilitates operation of aircraft 300 asdescribed herein. Accordingly, direct bearing joint 314 may be formedbetween fuselage skin 304 and any one or more main landing gear joints,and is not limited to being formed at joint 324 as described herein.

In the exemplary implementation, upon landing, the landing gear impartsa large compression punch load, as shown by arrow 328, into fuselage302. In the exemplary implementation, compressive load 328 isapproximately one million pounds of force, or higher. Alternatively,load 328 may be any amount of force that requires use of direct bearingjoint 314 as described herein. Compressive load 328 is transferred towheel well longeron 308 from a single location at forward main landinggear joint 324. Load 328 is then transferred from longeron 308 to skin304 through direct bearing joint 314 for distribution throughoutfuselage 302. As described above, to transfer large loads, conventionalmethods of load transfer requires the use an excessive amount of largefasteners spread over a large area. However, the use of such fastenerswould increase the weight, cost and complexity of the aircraft. Incontrast, direct bearing joint 314 reduces the size and number offasteners 310 required to couple skin 304 to longeron 308 for thedistribution of large compression load 328 as is described in furtherdetail below.

FIG. 5 is an enlarged view of a portion of aircraft fuselage 302outlined in box 5 shown in FIG. 4 illustrating direct bearing joint 314.In the exemplary implementation, joint 314 is positioned proximateforward main landing gear joint 324 and at least partially overlaps abulkhead 330 of plurality of frames 306 such that load 328 istransferred through wheel well longeron 308 and principal bulkhead 330to be distributed through skin 304. Alternatively or in combination,joint 314 may be positioned at any location on aircraft 300 where alarge compressive load transfer takes place between skin 304 and anotheraircraft structural component, such as the keel beam and/or the rearwing spar. Forming direct bearing joint 314 includes positioninglongeron 308 in intimate contact with skin 304 and coupling longeron 308and skin 304 together with fasteners 310. Longeron 308 includes ahorizontal portion 332 that is substantially perpendicular to load 328and a vertical portion 334 that is substantially parallel to load 328.Vertical portion 334 includes an extension flange 336 that extends fromvertical portion 334 in a direction substantially parallel to load 328.Extension flange 336 overlaps at least a portion of bulkhead 330 suchthat skin 304 is coupled between flange 336 and bulkhead 330 tofacilitate strengthening joint 314. Skin 304 includes a distal skin edge338 that contacts a portion of longeron 308 at direct bearing joint 314,but that is also positioned a distance away from longeron horizontalportion 332 such that a gap is formed between horizontal portion 332 andskin edge 338 on either side of direct bearing joint 314 along fuselage302.

In the exemplary implementation, direct bearing joint 314 includes afirst arcuate feature, or an arcuate cutout 340 from skin edge 338 thatis configured to receive a correspondingly shaped second arcuatefeature, or platform 342, on vertical portion 334 of longeron 308. Insuch a configuration, skin 304 and longeron 308 are in intimate contactalong arch-shaped direct bearing joint 314 such that concentratedcompression load 328 is distributed through skin 304 reducing stressconcentration, and enabling reaction of non-vertical load components. Inthe exemplary implementation, cutout 340 includes an arcuate cutout edge344 having a longitudinal width W along axis 312 and a varying height Hthat extends along vertical portion 334 between skin edge 338 and cutoutedge 344. The dimensions of height H and width W are optimized based onboth the size of both aircraft 300 and load 328 such that direct bearingjoint 314 best distributes load 328 through skin 304. In the exemplaryimplementation, cutout 340 is at least one of, but not limited to,elliptical and/or semi-circular, and/or rational B-spline.Alternatively, cutout 340 may be any shape, such as linear, thatfacilitates operation of direct bearing joint 314 as described herein.

Platform 342 of longeron 308 extends between inward vertical portion 334and bulkhead 330 and includes a height H and a width W that aresubstantially similar to that of cutout 340 such that platform 342provides an arcuate step that is configured to correspond to adjacentarcuate cutout 340. Platform height H and width W define a shape that iscomplementary to the arcuate shape of cutout 340 to facilitate loadtransfer therebetween. More specifically, platform 342 includes aconcave platform edge 346 that is directly abutted against convex cutoutedge 344 of cutout 340. The arcuate shape of direct bearing joint 314facilitates distributing load 328 over a larger area of cutout edge 344along width W than would be possible with a simply linear interface. Thearcuate shape of the direct bearing joint 314 also facilitates transferof loads with components that are not completely vertical.

FIG. 6 is a cross-sectional view of direct bearing joint 314 at alocation 6-6 (shown in FIG. 4) along aircraft fuselage 302. Similarly,FIG. 7 is a cross-sectional view of the interface between skin 304 andlongeron 308 at a location that is at a distal end of direct bearingjoint 314, such as location 7-7 (shown in FIG. 4) along fuselage 302that is aft of location 6-6. FIG. 6 illustrates a cross-sectional viewof direct bearing joint 314 at location 6-6 along a centerline ofprincipal bulkhead 330 where height H is at a maximum. In the exemplaryimplementation, skin 304 includes a thickness T₁ that is substantiallysimilar to a thickness T₂ of platform 342. Furthermore, extension flange336 includes a thickness T₃ such that a base 348 of vertical portion 334proximate horizontal portion 332 has a thickness T₄ equal to extensionflange thickness T₃ added to platform thickness T₂. Skin 304 alsoincludes an inner surface 350 that is coupled in a face-to-facerelationship with an outer surface 352 of principal bulkhead 330.Moreover, platform 342 includes an inner surface 354 that coupled in aface-to-face relationship with outer surface 352 such that innersurfaces 350 and 354 are substantially flush.

FIGS. 6 and 7 illustrate a continuous interface between skin 304 andlongeron 308 along direct bearing joint 314. Cutout edge 344 is incontinuous contact with platform edge 346 along direct bearing joint314, as shown in FIG. 6 at location 6-6, such that there is no gapbetween skin 304 and longeron 308. Such a continuous interfacefacilitates distributing punch compression load 328 through directbearing joint 314 into skin 304. Furthermore, should compression load328 contain a non-vertical component, the arcuate shape of directbearing joint 314 facilitates distributing such stress concentrationsmore evenly through skin 304. FIG. 7 is taken at an aft terminationpoint of direct bearing joint 314, and more specifically, of platform342. FIG. 7 illustrates that contact between skin edge 338 and longeron308 ends at the distal ends of direct bearing joint 314. Morespecifically, skin edge 338 is positioned a distance away from longeronhorizontal portion 332 such that a gap (best shown in FIG. 5) is definedtherebetween.

In the exemplary implementation, longeron 308, skin 304, and principalbulkhead 330 are coupled together with plurality of fasteners 310. Atlocation 6-6, fasteners 310 extend through extension flange 336, skin304, and principal bulkhead 330 and also through base 348, includingplatform 342, and bulkhead 330. At location 7-7, fasteners 310 extendsimply through vertical portion 334 and skin 304. Fasteners 310 may benecessary to secure longeron 308 and skin 304 together, although somefasteners 310 may not penetrate skin 304. However, because at least aportion of load 328 is directly transferred into skin 304 through directbearing joint 314, fasteners 310 carry only a portion of load 328.Accordingly, the number and size of fasteners 310 required at joint 314to carry load 328 is significantly reduced as compared to aircrafthaving a gap or a linear relationship between the skin and longeron.More specifically, the continuous interface between cutout 340 andplatform 342 and the plurality of fasteners 310 that couple skin 304 tolongeron 308 form two distinct load paths through which load 328 isdistributed. Furthermore, because direct bearing joint 314 does notcarry tension loads, fasteners 310 are configured to carry any tensionloads that may be imparted onto direct bearing joint 314. The tensionloads at direct bearing joint 314 are generally much smaller thancompression load 328, thus enabling them to be carried by the smallerand fewer fasteners 310.

FIG. 8 is a cross-sectional view of an alternative direct bearing joint400 at location 8-8 along aircraft fuselage 302 (shown in FIG. 4).Location 8-8 is the same location as the cross-section taken at location6-6 and shown in FIG. 6. Direct bearing joint 400 is substantiallysimilar to direct bearing joint 314 (shown in FIGS. 4-6) in operationand composition such that skin 304 is coupled to second aircraftstructure 206 (shown in FIG. 3), with the exception that direct bearingjoint 400 includes skin 304 coupled to a bearing plate 402 that iscoupled to a load bearing structural component of aircraft 300. As such,like components shown in FIG. 8 are labeled with like reference numbersused in FIGS. 4-6. Direct bearing joint 400 may be used in cases whereit may be easier to fabricate bearing plate 402 for coupling to atraditional (non-modified) longeron, keel beam, or wing rear spar frameinstead of modifying one or more of the metal aircraft structures toinclude platform 342 and extension flange 336. In such animplementation, plate 402 is fabricated from a metallic material, suchas, but not limited to titanium, aluminum, steel, or any combinationthereof. Alternatively, plate 402 may be fabricated from any suitablemetal or metal alloy that facilitates operation of joint 202 asdescribed herein.

As shown in FIG. 8, the structural component is a wheel well longeron404. However, the structural component may be any structural component,such as, but not limited to, a keel beam and/or a wing rear spar frame.In the implementation shown in FIG. 8, direct bearing joint 400 includesskin 304, bulkhead 330, wheel well longeron 404 having a horizontalportion 406 and a vertical portion 408, and bearing plate 402. Bearingplate 402 is coupled to vertical portion 408 and includes a distal plateedge 410 that is coupled in contact with horizontal portion 406. Bearingplate 402 also includes an extension flange 412 and an arcuate platform414 that are substantially similar to extension flange 336 and platform342, respectively. Similarly to direct bearing joint 314, cutout 340 isconfigured to receive correspondingly-shaped platform 414 such that skin304 and plate 402 are in intimate contact along arch-shaped directbearing joint 400 to facilitate distribution of concentrated compressionload 328 through skin 304 without using as many large fasteners as wouldbe required given a linear interface between skin edge 338 and plateedge 410. Alternatively, direct bearing joint 400 may have any shape,such as, but not limited to, a linear shape, that facilitates operationof joint 400 as described herein.

Platform 414 of bearing plate 402 extends between vertical portion 408and principal bulkhead 330 and includes a height H and a width W (notshown) that are substantially similar to that of cutout 340 such thatplatform 414 provides an arcuate step that is configured to correspondto adjacent arcuate cutout 340. More specifically, platform 414 includesa concave platform edge 416 that is directly abutted against convexcutout edge 344 of cutout 340. The arcuate shape of direct bearing joint400 facilitates distributing load 328 over a larger area of cutout edge344 along width W than would be possible with a simply linear interface.

In the implementation shown in FIG. 8, skin 304 and platform 414 includethickness T₁ and extension flange 412 includes a thickness T₅ that issubstantially equal to thickness T₃ (shown in FIG. 6) such that a base418 of plate 402 proximate horizontal portion 406 has a thickness T₆that is substantially similar to thickness T₄(shown in FIG. 6). Skininner surface 350 is coupled in a face-to-face relationship with outersurface 352 of bulkhead 330. Moreover, platform 414 includes an innersurface 420 that is coupled in a face-to-face relationship with outersurface 352 such that inner surfaces 350 and 420 are substantially flushwith each other.

Direct bearing joint 400 includes a continuous interface between skin304 and plate 402. More specifically, cutout edge 344 is in continuouscontact with platform edge 416 along direct bearing joint 400 such thatthere is no gap between skin 304 and plate 402 along joint 400. Such acontinuous interface facilitates distributing punch compression load 328through direct bearing joint 400 into skin 304. Furthermore, shouldcompression load 328 contain a non-vertical component, the arcuate shapeof direct bearing joint 400 facilitates distributing such stressconcentrations more evenly through skin 304.

In this implementation, longeron 308, skin 304, bulkhead 330, andbearing plate 402 are coupled together with plurality of fasteners 310.At location 8-8, fasteners 310 extend through extension flange 412, skin304, and bulkhead 330 and also through vertical portion 408, base 418,including platform 414, and bulkhead 330. Fasteners 310 may be necessaryto secure longeron 308 and skin 304 together, although some fasteners310 may not penetrate skin 304. However, because at least a portion ofload 328 is directly transferred into skin 304 through direct bearingjoint 400, fasteners 310 carry only a portion of load 328. Accordingly,the number and size of fasteners 310 required at joint 400 to carry load328 is significantly reduced as compared to aircraft having a gap or alinear relationship between the skin and longeron. More specifically,the continuous interface between cutout 340 and platform 414 and theplurality of fasteners 310 that couple skin 304 to plate 402 form twodistinct load paths through which load 328 is distributed compared toaircraft with only a single load path through the fasteners.

FIG. 9 is another alternative direct bearing joint 500 that may be usedwith aircraft fuselage 302 (shown in FIG. 4) and taken at location 9-9(shown in FIG. 4). FIG. 10 is a cross-sectional view of joint 500 takenat location 10-10. Direct bearing joint 500 is substantially similar todirect bearing joints 314 and 400 (shown in FIGS. 4-6 and 8) inoperation and composition such that skin 304 is coupled to secondaircraft structure 206 (shown in FIG. 3), with the exception that directbearing joint 500 includes skin 304 coupled to a keel chord extension502 rather than longeron 308. In this alternative implementation,aircraft 300 (shown in FIG. 4) includes a keel chord extension 502, akeel chord 504, and a keel joint 506 therebetween. Keel chord extension502 and keel chord 504 serve as a keel beam for aircraft 300 that isconfigured to transfer a compressive load 508 into skin 304. Aircraft300 also includes an aft wheel well chord 510 that is coupled to atleast a portion of both keel chord extension 502 and skin 304. Aft wheelwell chord 510 includes a horizontal portion 512 that is substantiallyparallel to load 508 and a vertical portion 514 that is substantiallyperpendicular to load 508. Horizontal portion 512 includes an extensionflange 516 that extends from horizontal portion 512 in a directionsubstantially parallel to load 508. Extension flange 516 overlaps atleast a portion of keel chord extension 502 such that skin 304 iscoupled between flange 516 and keel chord extension 502 to facilitatestrengthening direct bearing joint 500. Skin 304 includes a distal skinedge 518 that terminates proximate vertical portion 514 of aft wheelwell chord 510 at keel joint 506.

In this implementation, direct bearing joint 500 includes a cutout 520from skin edge 518 that is configured to receive a correspondinglyshaped platform 522 that extends from keel chord extension 502. Cutout520 and platform 522 are substantially similar in composition andfunction to cutout 340 and platform 342, respectively. In such aconfiguration, skin 304 and keel chord extension 502 are in intimatecontact along arch-shaped direct bearing joint 500 such thatconcentrated compression load 508 is distributed through skin 304,reducing stress concentration and enabling reaction of non-vertical loadcomponents. In this implementation, cutout 520 includes an arcuatecutout edge 524 having a width W and a varying length L that extendsalong horizontal portion 512 of aft wheel well chord 510 between skinedge 518 and cutout edge 524. The dimensions of length L and width W areoptimized based on both the size of both aircraft 300 and load 508 suchthat direct bearing joint 500 best distributes load 508 through skin304.

Platform 522 of keel chord extension 502 extends along horizontalportion 512 of aft wheel well chord 510 and includes a length L and awidth W that are substantially similar to that of cutout 520 such thatplatform 522 provides an arcuate step that is configured to correspondto adjacent arcuate cutout 520. Platform 522 length L and width W definea shape that is complementary to the arcuate shape of cutout 520 tofacilitate load transfer therebetween. More specifically, platform 522includes a concave platform edge 526 that is directly abutted againstconvex cutout edge 524 of cutout 520. The arcuate shape of directbearing joint 500 facilitates distributing load 508 over a larger areaof cutout platform edge 526 along width W than would be possible with asimply linear interface. The arcuate shape of direct bearing joint 500also facilitates transfer of loads with components that are orientedobliquely with respect to load 508.

Direct bearing joint 500 includes a continuous interface between skin304 and keel chord extension 502. More specifically, cutout edge 524 isin continuous contact with platform edge 526 along direct bearing joint500 such that there is no gap between skin 304 and keel chord extension502 along joint 500. Such a continuous interface facilitatesdistributing compression load 508 through direct bearing joint 500 intoskin 304. Furthermore, the arcuate shape of direct bearing joint 500facilitates distributing loads that contain a load component orientedobliquely with respect to load 508 more evenly through skin 304.

In this implementation, fasteners 310 may be necessary to couple aftwheel well chord 510, skin 304, and keel chord extension 502 together.However, some fasteners 310 may not penetrate skin 304 and extend onlythough aft wheel well chord 510 and keel chord extension 502, morespecifically, platform 522 of keel chord extension 502. Fasteners 310are configured to carry only a portion of load 508 because a portion ofload 508 is directly transferred into skin 304 from keel chord extension502. Accordingly, the number and size of fasteners 310 required at joint500 to carry load 508 is significantly reduced as compared to aircrafthaving a gap or a linear relationship between the skin and keel chordextension. More specifically, the continuous interface between cutout520 and platform 522 and the plurality of fasteners 310 that couple skin304 to keel chord extension 502 form two distinct load paths throughwhich load 508 is distributed compared to aircraft with only a singleload path through the fasteners.

The implementations described herein facilitate direct loading of acomposite skin that is coupled to a metal component. More specifically,direct loading of the composite skin is accomplished by forming a directbearing joint having an arcuate continuous interface between the skinand a load bearing structure. For example, the skin includes an arcuatecutout that directly contacts a correspondingly-shaped platform on aload bearing structure at a point where a large compressive load isapplied to the structure. Joints between composite and metal structuresare generally held together with fasteners. The methods described hereinfacilitate distributing large compressive loads by directly transferringat least a portion of the load into the skin such that a first load pathis formed at the joint between the skin and the structure and a secondload path is formed through the fasteners. Multiple loads paths reducethe number and size of the fasteners required to safely transfer thelarge loads into the composite skin, thus reducing the weight, cost, andassembly time of the structure having the direct bearing joint.

This written description uses examples to disclose variousimplementations, including the best mode, and also to enable any personskilled in the art to practice the various implementations, includingmaking and using any devices or systems and performing any incorporatedmethods. The patentable scope of the disclosure is defined by theclaims, and may include other examples that occur to those skilled inthe art. Such other examples are intended to be within the scope of theclaims if they have structural elements that do not differ from theliteral language of the claims, or if they include equivalent structuralelements with insubstantial differences from the literal language of theclaims.

What is claimed is:
 1. A method of forming a direct bearing joint, saidmethod comprising: providing a load bearing structure including a firststructural feature having an arcuate shape; providing a composite skinincluding a second structural feature having an arcuate shape; andcoupling the first structural feature to the second structural featuresuch that the first and second structural features are mated againsteach other to facilitate distributing compressive loads induced into theload bearing structure into the composite skin.
 2. The method inaccordance with claim 1, wherein providing a load bearing structureincluding a first structural feature further comprises providing a loadbearing structure including a convex platform edge, and whereinproviding a composite skin including a first structural feature furthercomprises providing a composite skin including a concave cutout edge. 3.The method in accordance with claim 2, wherein coupling the firststructural feature to the second structural feature further comprisesmating the convex platform edge in continuous direct contact against theconcave cutout edge.
 4. The method in accordance with claim 1, furthercomprising coupling the load bearing structure to the composite skinusing a plurality of fasteners such that a first load path is definedthrough the first and second structural features and a second load pathis defined through the plurality of fasteners.
 5. The method inaccordance with claim 1, wherein providing a load bearing structurefurther comprises providing one of a wheel well longeron, a keel beam,and a wing rear spar frame of an aircraft.
 6. The method in accordancewith claim 1, wherein providing a load bearing structure furthercomprises providing a bearing plate that is coupled to an airframecomponent selected from a group consisting of a wheel well longeron, akeel beam, and a wing rear spar frame of an aircraft.
 7. The method inaccordance with claim 1, wherein the load bearing structure includes astructural flange, the method further comprising coupling the compositeskin between the structural flange and a bulkhead.
 8. The method inaccordance with claim 7 further comprising: coupling a first innersurface of the second structural feature in a face-to-face relationshipwith an outer surface of the bulkhead; and coupling a second innersurface of the composite skin in a face-to-face relationship with theouter surface such that the first and second inner surfaces are flush.9. A direct bearing joint comprising: a load bearing structurecomprising a first structural feature that has an arcuate shape; and acomposite skin coupled to said load bearing structure, said compositeskin comprising a second structural feature that has an arcuate shape,said second structural feature configured to mate against said firststructural feature to facilitate distributing compressive loads inducedinto said load bearing structure into said composite skin.
 10. The jointin accordance with claim 9, wherein said first structural feature is aconvex platform and said second structural feature is a complementaryconcave cutout.
 11. The joint in accordance with claim 9, wherein saidfirst structural feature comprises a first edge, said second structuralfeature comprises a second edge that is shaped to mate against saidfirst edge.
 12. The joint in accordance with claim 9, wherein said loadbearing structure is one of a wheel well longeron, a keel beam, and awing rear spar frame of an aircraft.
 13. The joint in accordance withclaim 9, further comprising a plurality of fasteners configured tocouple said composite skin to said load bearing structure, such that afirst load path is defined through said first and second structuralfeatures, and a second load path is defined through said plurality offasteners.
 14. The joint in accordance with claim 9, wherein said loadbearing structure is a bearing plate coupled to one of a wheel welllongeron, a keel beam, and a wing rear spar frame of an aircraft. 15.The joint in accordance with claim 9, further comprising a bulkhead,said load bearing structure further comprises an extension flange, saidcomposite skin is coupled between said extension flange and saidbulkhead.
 16. The joint in accordance with claim 15, wherein said secondarcuate feature comprises a first inner surface coupled in a matingrelationship with an outer surface of said bulkhead, said composite skincomprises a second inner surface coupled in a mating relationship withsaid outer surface of said bulkhead such that said first and secondinner surfaces are substantially flush against each other.
 17. Anaircraft comprising a direct bearing joint, said aircraft comprising: aload bearing structure comprising a first structural feature that has anarcuate shape; a composite skin coupled to said load bearing structure,said composite skin comprising a second structural feature that has anarcuate shape, said second structural feature configured to mate againstsaid first structural feature to facilitate distributing compressiveloads induced into said load bearing structure into said composite skin.18. The aircraft in accordance with claim 17, wherein said firststructural feature comprises a first edge, said second structuralfeature comprises a second edge that is shaped to mate against saidfirst edge.
 19. The aircraft in accordance with claim 17, wherein saidload bearing structure is one of a wheel well longeron, a keel beam, anda wing rear spar frame.
 20. The aircraft in accordance with claim 17further comprising a plurality of fasteners configured to couple saidcomposite skin to said load bearing structure, such that a first loadpath is defined through said first and second structural features, and asecond load path is defined through said plurality of fasteners.